The present invention relates generally to both gas turbines and steam turbines, and more particularly to internal core cooling arrangements in turbine blades. Turbine blades are typically cooled to decrease both the bulk temperature and maximum temperature of the turbine blade material to support higher mechanical loads in the turbine blade that incorporates an improved aerodynamic performance configuration.
Steam turbines include, but are not limited, to steam turbine power generation equipment and shipboard steam turbine propulsion equipment. Gas turbines include, but are not limited to, gas turbine power generation equipment and gas turbine aircraft engines. An exemplary steam turbine typically contains a high-pressure turbine section, a low-pressure turbine section, or a combination of both, which is rotated by the steam flow. An exemplary gas turbine typically includes a core engine, having a high pressure compressor to compress the air flow entering the core engine, a combustor in which a mixture of fuel and the compressed air is burned to generate a propulsive gas flow, and a high pressure turbine which is rotated by the propulsive gas flow and which is connected by a larger diameter shaft to drive the high pressure compressor. A typical front fan gas turbine aircraft engine adds a low pressure turbine (located aft of the high pressure turbine) connected by a smaller diameter coaxial shaft to drive the front fan (located forward of the high pressure compressor) and to drive an optional low pressure compressor (located between the front fan and the high pressure compressor). The low-pressure compressor is sometimes called a booster compressor or simply a booster.
In the exemplary gas turbine, typically the fan and the high and low pressure compressors and turbines have gas turbine blades each including an airfoil portion attached to a shank portion. In the exemplary steam turbine, typically the high and low pressure turbine sections have steam turbine blades each including an airfoil portion attached to a shank portion. Rotor blades are gas or steam turbine blades attached to a rotating gas or steam turbine rotor discs, respectively. Stator vanes are gas turbine blades or steam turbine blades attached to a non-rotating gas or steam turbine stator casings, respectively. Typically, there are alternating circumferential rows of radially-outwardly extending rotor blades and radially-inwardly extending stator vanes. When present in the gas turbine configuration, at least one first and one last row of stator vanes (also called inlet and outlet guide vanes) typically have their radially-inward ends also attached to a non-rotating gas turbine stator casing. Counter rotating xe2x80x9cstatorxe2x80x9d vanes are also known in gas turbine designs. Conventional gas and steam turbine blade designs typically have airfoil portions that are made entirely of metal, such as titanium, or are made entirely of a composite. The all-metal blades, including costly wide-chord hollow blades, are heavier in weight, resulting in lower fuel performance and requiring sturdier blade attachments.
In a gas turbine aircraft application, the gas turbine blades that operate in the hot gas path are exposed to some of the highest temperatures in the gas turbine. Various design schemes have been pursed to increase the longevity of the blades in the hot gas path. By way of example and not limitation, these design schemes include blade coatings, and internal cooling of the blades.
In one common internal core cooling arrangement, a series of radial cooling holes extend through the entire turbine blade. The turbine blade is first manufactured as a solid blade. The solid blade is then drilled using Electro-Chemical Machining (ECM) or Shaped-Tube Electro-Chemical Machining (STEM), to create a plurality of through holes from about a blade root to about a blade tip. The radial cooling holes in axially long blades can be difficult to machine, sometimes requiring drilling from both ends of the blade. The blade with the radial cooling holes tends to have more mass than is desired. The extra mass can be problematic during thermal transients as the interior surfaces and the exterior surfaces of the blade do not respond at the same rate to the thermal transient, which results in thermal stresses. Moreover, the use of radial cooling holes is generally not possible in the leading and trailing edges of the blades, due to the three dimensional curvature of the blade. Alternatively, the need to locate the radial cooling holes forces a compromise of the aerodynamics to accommodate straight holes. One design alternative to the radial cooling holes is to bleed cooling flow to form a film-cooling layer over the blade.
The coolant for the internal cooling of the blades typically comes from a cooler temperature part of the gas turbine or from a separate source of cooling. The coolant is typically either an air-based coolant or a steam-based coolant. The air-based coolant is typically bled either from the compressor section or from a post-compressor region that surrounds the combustion section that is operating at a cooler temperature than the turbine blades and blade covers of concern. The air-based coolant is alternately supplied from a separate off-machine located air supply system. The steam-based coolant is typically supplied from a turbine section that is operating at a cooler temperature than the turbine blades of concern or the steam-based coolant can be supplied from an independent steam supply (i.e. other steam system or auxiliary boiler). However, providing the air-based coolant to internally cool the turbine blades represents internal work to the gas turbine that reduces the net output power of the gas turbine. Additionally, the issues related to directing the flow of the air-based coolant to the areas of highest heat load in the turbine blade has created the desire to improve the internal cooling of the blades even further.
Accordingly, there is a need for an improved turbine blade. What is needed is a turbine blade core cooling apparatus that allows more aggressively shaped aerodynamic blade configurations, promotes lighter blade internal construction, maintains the structural support of the turbine blade, delivers higher cooling effectiveness, and lowers sensitivity to wall thickness variations by placing cooling air very near all external surfaces of the turbine blade. What is also needed is an internal cooling scheme that satisfies the turbine blade cooling requirements with less impact on the turbine net output.
The present invention provides a cooling apparatus for cooling a turbine blade. The cooling apparatus comprises a pressure side plate comprising a plurality of pressure ribs, a suction side plate comprising a plurality of suction ribs, and a plurality of flow redirection areas. The pressure side plate is disposed over the suction side a plate, where the pressure ribs are disposed at a first angle with respect to a blade span reference line and the suction ribs are disposed at a second angle with respect to the blade span reference line, to form the flow redirection areas.
The present invention provides a method of fabricating the cooling apparatus for the turbine blade comprising aligning the pressure side plate and the suction side plate to form the plurality of flow redirection areas between the pressure side plate and the suction side plate.